This invention relates generally to the field of gas turbine engines, and more particularly to an internally cooled hybrid ceramic matrix composite vane.
Gas turbine engines are known to include a compressor section for supplying a flow of compressed combustion air, a combustor section for burning a fuel in the compressed combustion air, and a turbine section for extracting thermal energy from the combustion air and converting that energy into mechanical energy in the form of a shaft rotation. Many parts of the combustor section and turbine section are exposed directly to the hot combustion gasses, for example the combustor, the transition duct between the combustor and the turbine section, and the turbine stationary vanes, rotating blades and surrounding ring segments.
It is also known that increasing the firing temperature of the combustion gas may increase the power and efficiency of a combustion turbine. Modern, high efficiency combustion turbines have firing temperatures in excess of 1,600xc2x0 C., which is well in excess of the safe operating temperature of the structural materials used to fabricate the hot gas flow path components. Accordingly, several methods have been developed to permit operation of these materials in this environment. These include film cooling, backside cooling and thermal barrier coatings.
Film cooling involves the delivery of a film of cooling fluid, such as compressed air extracted from the compressor section, between the structural component and the flow of hot combustion gasses. The film of cooling fluid may be provided from a bleed flow from the compressor through holes formed in the surface of the component to be cooled. Film cooling systems are generally very effective in cooling a component, however they may significantly reduce the efficiency of the machine. Energy is needed to compress the cooling fluid, a decrease in combustion gas temperature is induced by the addition of the relatively cold fluid, and disturbance may be created in the smooth flow of air over an airfoil component such as a blade or vane.
Backside cooling generally involves the passage of a cooling fluid over a backside of a component that has a front side exposed to the hot combustion gasses. The cooling fluid in backside cooling schemes may be compressed air that has been extracted from the compressor or steam that is available from other fluid loops in a combustion turbine power plant. Backside cooling does not affect the exhaust gas composition or the flow of air over an airfoil component, it does not dilute the hot combustion air with colder fluid, and it can generally be supplied at a lower pressure than would be needed for film cooling. However, backside cooling creates a temperature gradient across the thickness of the cooled wall, and thus becomes decreasingly effective as the thickness of the component wall increases and as the thermal conductivity of the material decreases.
Insulation materials such as ceramic thermal barrier coatings (TBC""s) have been developed for protecting temperature-limited components. While TBC""s are generally effective in affording protection for the current generation of combustion turbine machines, they may be limited in their ability to protect underlying metal components as the required firing temperatures for next-generation turbines continue to rise.
Ceramic matrix composite (CMC) materials offer the potential for higher operating temperatures than do metal alloy materials due to the inherent nature of ceramic materials. This capability may be translated into a reduced cooling requirement that, in turn, may result in higher power, greater efficiency, and/or reduced emissions from the machine. However, CMC materials generally are not as strong as metal, and therefore the required cross-section for a particular application may be relatively thick. Due to the low coefficient of thermal conductivity of CMC materials and the relatively thick cross-section necessary for many applications, backside closed-loop cooling is generally ineffective as a cooling technique for protecting these materials in combustion turbine applications. Accordingly, high temperature insulation for ceramic matrix composites has been described in U.S. Pat. No. 6,197,424 B1, which issued on Mar. 6, 2001, and is commonly assigned with the present invention. That patent describes an oxide-based insulation system for a ceramic matrix composite substrate that is dimensionally and chemically stable at a temperature of approximately 1600xc2x0 C. That patent also describes a stationary vane for a gas turbine engine formed from such an insulated CMC material. A similar gas turbine vane 10 is illustrated in FIG. 1 as including an inner wall 12 and stiffening ribs 14 formed of CMC material covered by an overlying layer of insulation 16. Backside cooling of the inner wall 12 is achieved by convection cooling, e.g. via direct impingement through supply baffles (not shown) situated in the interior chambers 18 using air directed from the compressor section of the engine.
If baffles or other means are used to direct a flow of cooling fluid throughout the airfoil member for backside cooling and/or film cooling, the cooling fluid is typically maintained at a pressure that is in excess of the pressure of the combustion gasses on the outside of the airfoil so that any failure of the pressure boundary will not result in the leakage of the hot combustion gas into the vane. Such cooling passages must generally have a complex geometry in order to provide a precise amount of cooling in particular locations to ensure an adequate degree of cooling without over-cooling of the component. It is generally very difficult to form such complex cooling passages in a ceramic matrix composite component. Alternatively, large central chambers 18 as illustrated in FIG. 1 may be used with appropriate baffling to create impingement of the cooling fluid onto the backside of the surface to be cooled. Such large chambers create an internal pressure force that can result in the undesirable ballooning of the airfoil structure due to the internal pressure of the cooling fluid applied to the large internal surface area of the passage 18. Furthermore, the geometry of FIG. 1 is also limited by stress concentrations at the intersection of the stiffening ribs 14 and the inner wall 12.
Even higher operating temperatures are envisioned for future generations of combustion turbine machines. Accordingly, further improvements in the design of ceramic matrix composite airfoils and the cooling of such airfoils are needed.
Accordingly, a hybrid turbine component is described herein as including a CMC airfoil member defining a core region and a core member bonded to the airfoil member within the core region. The core member includes cooling channels for the passage of a cooling fluid for removing heat from the CMC material through the bond. The cooling passage may be formed as a groove on an outside surface of the core member, thereby providing both convective and conductive cooling of the CMC member. By bonding the core member to at least 30% of the inside area of the CMC airfoil member, the internal stress caused by the cooling fluid pressure is reduced. An insulating material may be deposited over the CMC airfoil member to reduce the cooling flow requirements. The materials properties of the various components are selected to minimize the stresses in the system.